Cooled cooling air system for a gas turbine engine

ABSTRACT

A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section via a primary flowpath and a turbine section fluidly connected to the combustor via the primary flowpath. Also included is a cascading cooling system having a first inlet connected to a first compressor bleed, a second inlet connected to a second compressor bleed downstream of the first compressor bleed, and a third inlet connected to a third compressor bleed downstream of the second compressor bleed. The cascading cooling system includes at least one heat exchanger configured to incrementally generate cooling air for at least one of an aft compressor stage and a foremost turbine stage relative to fluid flow through the turbine section.

TECHNICAL FIELD

The present disclosure relates generally to gas turbine engines, andmore specifically to a cooled cooling air system for a gas turbineengine.

BACKGROUND

Gas turbine engines, such as those utilized on commercial and militaryaircraft, include a compressor section that draws in air, a combustorsection that mixes the compressed air with a fuel, and ignites themixture, and a turbine section across which the results of thecombustion are expanded. The expansion across the turbine section drivesthe turbine section to rotate, which in turn drives rotation of thecompressor.

In some example engines, this configuration results in excess heat atthe aft stages of the compressor section and in the turbine section. Inorder to prevent the excess heat from damaging engine components, orreducing the lifecycle of engine components, portions of the compressorsection and the turbine section are actively cooled using cooled coolingair.

SUMMARY OF THE INVENTION

In one exemplary embodiment a gas turbine engine includes a compressorsection, a combustor fluidly connected to the compressor section via aprimary flowpath, a turbine section fluidly connected to the combustorvia the primary flowpath, and a cascading cooling system having a firstinlet connected to a first compressor bleed, a second inlet connected toa second compressor bleed downstream of the first compressor bleed, anda third inlet connected to a third compressor bleed downstream of thesecond compressor bleed. The cascading cooling system includes at leastone heat exchanger configured to incrementally generate cooling air forat least one of an aft compressor stage and a foremost turbine stagerelative to fluid flow through the turbine section.

In another exemplary embodiment of the above described gas turbineengine, the third compressor bleed is at an outlet of the compressorsection.

In another exemplary embodiment of any of the above described gasturbine engines, the at least one heat exchanger includes a first heatexchanger, a second heat exchanger in series with the first heatexchanger, and a third heat exchanger in series with the second heatexchanger.

In another exemplary embodiment of any of the above described gasturbine engines, a heat sink input of the second heat exchanger is acooled flow output of the first heat exchanger and originates at thefirst inlet, and wherein a heat sink input of the third heat exchangeris a cooled flow output of the second heat exchanger and originates atthe second inlet.

In another exemplary embodiment of any of the above described gasturbine engines, a cooled flow output of the third heat exchanger isprovided to an aft most compressor stage as a cooled cooling flow andoriginates at the third inlet.

In another exemplary embodiment of any of the above described gasturbine engines, the at least one heat exchanger includes a parallelheat exchanger having at least a first heat sink input, a first cooledflow input and a second cooled flow input, and wherein fluid passingthrough the first cooled flow input is simultaneously cooled by thefirst heat sink input and cools the second cooled flow input.

In another exemplary embodiment of any of the above described gasturbine engines, the at least one heat exchanger further includes athird cooled flow input, and fluid passing through the second cooledflow input is simultaneously cooled by the first cooled flow input andcools the third cooled flow input.

In another exemplary embodiment of any of the above described gasturbine engines, the third cooled flow input is returned to an aft mostcompressor stage as a cooled cooling flow.

In another exemplary embodiment of any of the above described gasturbine engines, further includes an auxiliary compressor system havingan auxiliary turbine connected to an auxiliary compressor such thatrotation of the auxiliary turbine drives rotation of the auxiliarycompressor, and wherein an input of the auxiliary turbine is connectedto a heat sink outlet of a third heat exchanger such that cooling airoriginating at the second compressor bleed is compressed in theauxiliary compressor.

In another exemplary embodiment of any of the above described gasturbine engines, an output of the auxiliary compressor system isprovided to a foremost stage of the turbine section as cooled coolingair.

In another exemplary embodiment of any of the above described gasturbine engines, an input of the auxiliary turbine is at least one of anoutput heat sink air of a second heat exchanger, and an output heat sinkair of a fourth heat exchanger, and wherein an output of the auxiliaryturbine is returned to a compressor section inlet.

In another exemplary embodiment of any of the above described gasturbine engines, the input of the auxiliary turbine is a combination ofa heat sink output of the second heat exchanger, and a heat sink outputof the fourth heat exchanger, and wherein the combination is controlledby a modulation valve.

In another exemplary embodiment of any of the above described gasturbine engines, fluid flow through the auxiliary turbine is at leastpartially controlled by a modulation valve downstream of an auxiliaryturbine outlet, and wherein fluid flow to the inlet of the auxiliarycompressor is at least partially controlled by a modulation valveconnecting the heat sink outlet of the third heat exchanger to a secondor later stage of the turbine section.

In another exemplary embodiment of any of the above described gasturbine engines, each of the modulation valves is controlled by at leastone controller.

In another exemplary embodiment of any of the above described gasturbine engines, the gas turbine engine is a geared turbofan engine.

In another exemplary embodiment of any of the above described gasturbine engines, the gas turbine engine is a multiple bypass flowengine.

An exemplary method for generating cooled cooling air in a gas turbineengine includes providing air from a plurality of compressor bleeds asinputs to a cascading heat exchanger, incrementally cooling air receivedvia the inputs such that a first of the inputs is a heat sink for asecond of the inputs, and the second of the inputs is a heat sink for athird of the inputs, and actively cooling at least one of an aft mostcompressor stage and a fore most turbine stage using cooled cooling airoutput from the cascading heat exchanger.

In a further example of the above described exemplary method forgenerating cooled cooling air in a gas turbine engine, incrementallycooling air received via the inputs includes cooling the air using aplurality of serially arranged heat exchangers, each of the seriallyarranged heat exchangers increasing a pressure of cooled output airrelative to a serially previous heat exchanger.

In a further example of any of the above described exemplary methods forgenerating cooled cooling air in a gas turbine engine, incrementallycooling air received via the inputs includes cooling the air using atleast one parallel heat exchanger configured such that air from at leastone of the inputs simultaneously cools an adjacent air flow and iscooled by another adjacent airflow.

In one exemplary embodiment a gas turbine engine includes a compressorsection, a combustor fluidly connected to the compressor section via aprimary flowpath, a turbine section fluidly connected to the combustorvia the primary flowpath, and a cascading cooling system having aplurality of inlets, each of the inlets connected to one of a pluralityof compressor bleeds. The cascading cooling system includes at least oneheat exchanger configured to incrementally generate a cooled cooling airacross a plurality of stages, each of the stages having approximatelythe same pressure differential as each other of the stages.

These and other features of the present invention can be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates a first exemplary gas turbine engine.

FIG. 2 schematically illustrates a second exemplary gas turbine engine.

FIG. 3 schematically illustrates a first example cooled cooling airsystem for a gas turbine engine.

FIG. 4 schematically illustrates the first example cooled cooling airsystem for a gas turbine engine with the addition of an auxiliarygenerator.

FIG. 5 schematically illustrates a second example cooled cooling airsystem for a gas turbine engine.

FIG. 6 schematically illustrates a third example cooled cooling airsystem for a gas turbine engine.

FIG. 7 schematically illustrates a fourth example cooled cooling airsystem for a gas turbine engine.

DETAILED DESCRIPTION OF AN EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

With regards to engines for military applications, there has recentlybeen provision of multiple bypass flow engines. An example multiplebypass flow engine is shown schematically in FIG. 2. A first stage fan180 delivers air into an outer housing 181. The outer housing 181defines an outer bypass duct 182 outwardly of an inner housing 183. Theouter bypass duct 182 is alternatively referred to as a “third stream”,and air from the outer bypass duct 182 is referred to as third streamair. A second stage fan 184 delivers air downstream of the first stagefan 180 into an inner bypass duct 186. The inner bypass duct 186 isdefined between an inner periphery of the inner housing 183 and an outerperiphery of a core housing 187. Core housing 187 defines a radiallyinner extent of inner bypass duct 186. Controls 196 and 198 are shownschematically. In one example, the controls are nozzles which controlthe flow of air through the bypass flow ducts 182 and 186.

The first stage fan 180 delivers air inwardly of outer housing 181 andinwardly of inner housing 183. A second stage fan 184 delivers airinwardly of inner housing 183, but does not deliver air inwardly of theouter housing 181.

A core engine inlet 188 receives air downstream of the second stage fan184. That air passes into a compressor 190, a combustor 192 and aturbine 194. It should be understood that the compressor 190 may includemultiple rotors and the turbine 194 may also comprise multiple rotors.The turbine rotors drive the compressor 190 and the fan stages 180 and184.

In some examples, the aft stages of the compressor section 24, 190 andthe turbine section 28, 194 of a given engine 20 are actively cooledusing cooled cooling air sourced at one or more compressor bleeds. Thecooling air bleed from the compressor section 24, 190 is cooled in aheat exchanger using air from the bypass flowpath B as a heat sink, inthe example of FIG. 1, or using air from the third stream 182, in theexample of FIG. 2. A fraction of air is bled from the compressor section24, 190 diffuser and cooled in a heat exchanger and is returned to thecompressor section 24, 190 and/or the turbine section 28, 194.

When the air is fully cooled in a single step, the pressure differentialbetween the cold side of the heat exchanger (the bypass flow path B orthe third stream 182) and the compressor bleed can be excessively high.Further, the magnitude of cooling required in a single cooling stepincurs a large thermal gradient across the heat exchanger. The highpressure differential and large thermal gradients places high stresseson the heat exchanger in the cooling step. The high stresses in turn,reduce the life cycle of the heat exchanger, require the heat exchangerto be excessively large and heavy to accommodate the higher stresses, orboth.

In order to minimize the weight requirements and the thermal stresses onthe heat exchanger, an example gas turbine engine 200, schematicallyillustrated in FIG. 3, utilizes a cascaded heat exchanger configurationto provide cooled cooling air to the segments of the engine 200 beingactively cooled. The example gas turbine engine 200 includes acompressor section 210 having multiple stages 212, a combustor section220, and a turbine section 230 having multiple stages 232.

A mid compressor bleed 240 withdraws a first bleed air 241 from acompressor flowpath between compressor stages 212 at approximatelymidway through the compressor section 210. The first bleed air 241 isprovided to a first heat exchanger 250. Also provided to the heatexchanger 250 is cold air 260 from the bypass flowpath (in the case of ageared turbofan engine) or the third stream (in the case of a multiplebypass flow engine). The cold air 260 is used as a heat sink in thefirst heat exchanger 250 to cool the first bleed air 241.

A second compressor bleed 242, aft of the mid compressor bleed 240,withdraws a second bleed air 243 from between compressor stages 212. Thesecond bleed air 243 is provided to a second heat exchanger 251. Alsoprovided to the second heat exchanger 251 is the cooled air output ofthe first heat exchanger 250. The cooled air output of the first heatexchanger 250 is the first bleed air 241. The first bleed air 241 actsas a heat sink for the second bleed air 243 in the second heat exchanger251.

A third compressor bleed 244 is positioned at an aft end of thecompressor section 210. The third compressor bleed 244 removes a thirdbleed air 245 from the compressor flowpath, and provides the third bleedair 245 to a third heat exchanger 252. Also provided to the third heatexchanger 252 is the cooled air output of the second heat exchanger 251.The cooled air output of the second heat exchanger 251 is the secondbleed air 243. The second bleed air 243 acts as a heat sink for thethird bleed air 245 in the third heat exchanger 252. Once cooled in thethird heat exchanger, the third bleed air is returned to the aftmoststage 212 of the compressor section 210 and actively cools the aftmoststage 212. Because the third bleed air 245 is pulled from a thirdcompressor bleed 244 at the aftmost stage, the pressure differencebetween the third bleed air 245 and the compressor flowpath at the pointwhere the third bleed air 245 is returned to the compressor flowpath isminimal.

The heat sink air from the second heat exchanger 251 is provided to afourth heat exchanger 253. As with the first heat exchanger 250, thefourth heat exchanger 253 is provided cold air 260 from the bypassflowpath (in the case of a geared turbofan engine) or the third stream(in the case of a multiple bypass flow engine). The cold air 260 coolsthe heat sink air from the second heat exchanger 251.

The heat sink air in the first heat exchanger 250 and the fourth heatexchanger 253 is drawn from the same source. As a result, in someconfigurations, the first heat exchanger 250 and the fourth heatexchanger 253 can be combined into a single heat exchanger 254. Theoutput heat sink air from both the first heat exchanger 250 and thefourth heat exchanger 253, or the combined heat exchanger 254, isreturned to the bypass flowpath or the third stream.

Also included within the gas turbine engine 200 is an auxiliarycompressor system 270. The auxiliary compressor system 270 includes anauxiliary turbine 272 and an auxiliary compressor 274. The cooled airoutput of the fourth heat exchanger 253 (the first bleed air 241) isprovided to the turbine 272, and expanded across the auxiliary turbine272. The expansion across the turbine drives rotation of the auxiliarycompressor 274. The auxiliary compressor 274, in turn, receives andcompresses at least a portion of the second bleed air 243 after thesecond bleed air 243 has been utilized as a heat sink in the third heatexchanger 252.

An auxiliary compressor output 276 is provided as cooled cooling air tothe first turbine stage 232. In some example cooling systems, a portionof the second bleed air 243 is provided to a second or later stage 232of the turbine section 230 and provides active cooling. The portion isremoved from the second bleed air 243 prior to the second bleed air 243being provided to the auxiliary compressor 274 at a branch 278. A valve280 controls the amount of air removed from the second bleed air 243prior to being provided to the auxiliary compressor 274. The valve 280is connected to and controlled by an engine controller 282, or any othersimilar control device.

In some examples, the flow of first bleed air 241 through the auxiliaryturbine 272 is controlled via a second valve 284. The second valve 284is also controlled via the engine controller 282. Expanded air 273 isoutput from the auxiliary turbine 272 and returned to an inlet of thecompressor section 210, where it is ingested and recompressed throughthe compressor section 210.

In some examples, the gas turbine engine 200 can further include anauxiliary generator 290. An exemplary gas turbine engine 200 with theadditional inclusion of an auxiliary generator 290 is illustrated inFIG. 4. The auxiliary generator 290 is connected to the auxiliarycompressor 274, and is driven to rotate by the rotation of the auxiliarycompressor 274. The rotation of the auxiliary generator 290 generateselectrical power according to known electricity generation principles.The auxiliary generator 290 can be connected to aircraft electricalsystems and provide electrical power to onboard electronic systems.

With continued reference to FIGS. 3 and 4, the pressure differentialbetween the first bleed air 241 and the heat sink air of the first heatexchanger 250 is relatively small. The heat sink configuration iscascaded upward in pressure, while maintaining approximately the samerelatively small pressure differential, such that the cooled air in eachheat exchanger 250, 251, 252, 253 is cooled by a heat sink air with arelatively small pressure differential. In some examples, the pressuredifferential in each heat exchanger 250, 251, 252, 253 is approximatelythe same. This upward cascade in cooling and pressure is referred to asa cascaded cooled cooling air system.

With continued reference to FIGS. 3 and 4, FIG. 5 illustrates analternate heat exchanger 250 arrangement for providing cascaded cooledcooling air to cool an aft stage of a compressor 310 and at least afirst stage of a turbine 330. As with the example of FIGS. 3 and 4, theexample gas turbine engine 300 of FIG. 5 includes a compressor section310 having multiple stages 312, a combustor section 320, and a turbinesection 330 having multiple stages 332. In place of the first, second,third and fourth heat exchangers 250, 251, 252, 253 of FIGS. 3 and 4,the example gas turbine engine of FIG. 5 utilizes a single parallel heatexchanger 350. The parallel heat exchanger 350 utilizes parallel flows,where a single flow acts as a heat sink to one adjacent flow and isactively cooled by another adjacent flow.

As with the examples of FIGS. 3 and 4, the example of FIG. 5 utilizes afirst compressor bleed 340 at an approximate mid-point in the compressor310, a second compressor bleed 341 aft of the first compressor bleed340, and a third compressor bleed aft of the second compressor bleed341. The cooled cooling air for each area in the example of FIG. 5 isgenerated in the same manner as described above with regards to theexamples of FIGS. 3 and 4. Similarly, an auxiliary compressor system 370is connected to the outputs of the parallel heat exchanger 350 in theexample of FIG. 5.

The heating and cooling flows through the parallel heat exchanger 350are cooled via identical heat sinks as the sequential heat exchangers250, 251, 252, 253 described above with regards to FIGS. 3 and 4, andprovide for the same heating and cooling generation. One of skill in theart, having the benefit of this disclosure, will understand that anynumber of sequential heat exchangers can be combined into at least oneparallel heat exchanger, provided the required heat duties sustain apositive temperature gradient between each adjacent pair of fluids..Further, the number of sequential heat exchangers and parallel heatexchangers to be used in a given example can be determined based on theavailable volume, weight limitations, and cooling requirements of anygiven engine.

With continued reference to FIGS. 3 and 4, FIG. 6 illustrates anotherexample configuration of a cascaded heat exchanger configuration forgenerating cooled cooling air in a gas turbine engine 400. The exampleof FIG. 5 utilizes a first compressor bleed 440 at an approximatemid-point in the compressor 410, a second compressor bleed 441 aft ofthe first compressor bleed 440, and a third compressor bleed 442 aft ofthe second compressor bleed 441. As with the examples of FIGS. 3 and 4,a first heat exchanger 450 and a fourth heat exchanger 453 utilize airfrom the bypass duct or the third stream as a heat sink. The first heatexchanger 450 receives air from the first compressor bleed 440 and coolsthe air using the heat sink air.

The cooled first bleed air 441 is provided as a heat sink in a parallelheat exchanger 451. The parallel heat exchanger 451 receives the secondbleed air 443 and the third bleed air 444. The second bleed air 443 issimultaneously cooled by the first bleed air 441 and cools the thirdbleed air 444. Each of the outputs of the parallel heat exchanger 451 isprovided to an auxiliary compressor system 470. The remainder of thefeatures and connections are substantially similar to the featuresillustrated in FIGS. 3 and 4 and described above.

With continued reference to FIGS. 3 and 4, FIG. 7 illustrates amodification that can be applied to the gas turbine engine 200 in anexample gas turbine engine 600. The heat exchangers 650, 651, 652, 653are arranged in the same sequential cascading manner as in the exampleof FIG. 3. In addition to the heat sink flows, and cooled air flows ofFIGS. 3 and 4, the example of FIG. 7 includes multiple diverter valves691, 692. The diverter valves 691, 692 enable an increased powerextraction from the auxiliary turbine 672. The fourth heat exchanger 653can be partially or wholly bypassed such that an inlet temperature ofthe auxiliary turbine 672 is higher. Higher inlet temperatures increasethe power output from the auxiliary turbine utilized to drive theauxiliary compressor 674 and an attached generator 690.

Further, the second diverter valve 691 shunts auxiliary turbinedischarge from the auxiliary turbine 672 to the fan bypass stream (inthe case of an engine according to FIG. 1) or the third stream (in thecase of an engine according to FIG. 2). The fan bypass stream or thethird stream is at a lower pressure than the inlet to the corecompressor. The expansion across the auxiliary turbine 672 is increasedto expand the cooled air to match the inlet pressure of the bypass flowor the third duct flow, allowing the auxiliary turbine to extract morework from the cooled air.

With reference to FIGS. 3-7, the illustrated heat exchangers and cooledcooling air systems can be located in any physical location within a gasturbine engine, subject to design restraints. By way of example, each ofthe heat exchangers and the auxiliary compressor system can bepositioned in a core cowling that surrounds the compressor, combustor,and turbine sections of the gas turbine engine. In alternative examples,the heat exchangers, auxiliary compressor system, generator, or anyother elements, can be disposed in any other location within the gasturbine engine.

It is further understood that any of the above described concepts can beused alone or in combination with any or all of the other abovedescribed concepts. Although an embodiment of this invention has beendisclosed, a worker of ordinary skill in this art would recognize thatcertain modifications would come within the scope of this invention. Forthat reason, the following claims should be studied to determine thetrue scope and content of this invention.

1. A gas turbine engine comprising: a compressor section; a combustorfluidly connected to the compressor section via a primary flowpath; aturbine section fluidly connected to the combustor via the primaryflowpath; a cascading cooling system having a first inlet connected to afirst compressor bleed, a second inlet connected to a second compressorbleed downstream of the first compressor bleed, and a third inletconnected to a third compressor bleed downstream of the secondcompressor bleed; the cascading cooling system including at least oneheat exchanger configured to incrementally generate cooling air for atleast one of an aft compressor stage and a foremost turbine stagerelative to fluid flow through the turbine section.
 2. The gas turbineengine of claim 1, wherein the third compressor bleed is at an outlet ofthe compressor section.
 3. The gas turbine engine of claim 1, whereinthe at least one heat exchanger includes a first heat exchanger, asecond heat exchanger in series with the first heat exchanger, and athird heat exchanger in series with the second heat exchanger.
 4. Thegas turbine engine of claim 3, wherein a heat sink input of the secondheat exchanger is a cooled flow output of the first heat exchanger andoriginates at the first inlet, and wherein a heat sink input of thethird heat exchanger is a cooled flow output of the second heatexchanger and originates at the second inlet.
 5. The gas turbine engineof claim 4, wherein a cooled flow output of the third heat exchanger isprovided to an aft most compressor stage as a cooled cooling flow andoriginates at the third inlet.
 6. The gas turbine engine of claim 1,wherein the at least one heat exchanger includes a parallel heatexchanger having at least a first heat sink input, a first cooled flowinput and a second cooled flow input, and wherein fluid passing throughthe first cooled flow input is simultaneously cooled by said first heatsink input and cools said second cooled flow input.
 7. The gas turbineengine of claim 5, wherein the at least one heat exchanger furtherincludes a third cooled flow input, and fluid passing through the secondcooled flow input is simultaneously cooled by said first cooled flowinput and cools said third cooled flow input.
 8. The gas turbine engineof claim 6, wherein said third cooled flow input is returned to an aftmost compressor stage as a cooled cooling flow.
 9. The gas turbineengine of claim 1 further comprising an auxiliary compressor systemhaving an auxiliary turbine connected to an auxiliary compressor suchthat rotation of the auxiliary turbine drives rotation of the auxiliarycompressor, and wherein an input of the auxiliary turbine is connectedto a heat sink outlet of a third heat exchanger such that cooling airoriginating at the second compressor bleed is compressed in theauxiliary compressor.
 10. The gas turbine engine of claim 9, wherein anoutput of the auxiliary compressor system is provided to a foremoststage of the turbine section as cooled cooling air.
 11. The gas turbineengine of claim 9, wherein an input of the auxiliary turbine is at leastone of an output heat sink air of a second heat exchanger, and an outputheat sink air of a fourth heat exchanger, and wherein an output of theauxiliary turbine is returned to a compressor section inlet.
 12. The gasturbine engine of claim 11, wherein the input of the auxiliary turbineis a combination of a heat sink output of said second heat exchanger,and a heat sink output of said fourth heat exchanger, and wherein thecombination is controlled by a modulation valve.
 13. The gas turbineengine of claim 9, wherein fluid flow through said auxiliary turbine isat least partially controlled by a modulation valve downstream of anauxiliary turbine outlet, and wherein fluid flow to said inlet of saidauxiliary compressor is at least partially controlled by a modulationvalve connecting the heat sink outlet of said third heat exchanger to asecond or later stage of said turbine section.
 14. The gas turbineengine of claim 13, wherein each of said modulation valves is controlledby at least one controller.
 15. The gas turbine engine of claim 1,wherein the gas turbine engine is a geared turbofan engine.
 16. The gasturbine engine of claim 1, wherein the gas turbine engine is a multiplebypass flow engine.
 17. A method for generating cooled cooling air in agas turbine engine comprising: providing air from a plurality ofcompressor bleeds as inputs to a cascading heat exchanger; incrementallycooling air received via said inputs such that a first of said inputs isa heat sink for a second of said inputs, and the second of said inputsis a heat sink for a third of said inputs; and actively cooling at leastone of an aft most compressor stage and a fore most turbine stage usingcooled cooling air output from said cascading heat exchanger.
 18. Themethod of claim 17, wherein incrementally cooling air received via saidinputs comprises cooling said air using a plurality of serially arrangedheat exchangers, each of the serially arranged heat exchangersincreasing a pressure of cooled output air relative to a seriallyprevious heat exchanger.
 19. The method of claim 17, whereinincrementally cooling air received via said inputs comprises coolingsaid air using at least one parallel heat exchanger configured such thatair from at least one of said inputs simultaneously cools an adjacentair flow and is cooled by another adjacent airflow.
 20. A gas turbineengine comprising: a compressor section; a combustor fluidly connectedto the compressor section via a primary flowpath; a turbine sectionfluidly connected to the combustor via the primary flowpath; a cascadingcooling system having a plurality of inlets, each of said inletsconnected to one of a plurality of compressor bleeds; the cascadingcooling system including at least one heat exchanger configured toincrementally generate a cooled cooling air across a plurality ofstages, each of said stages having approximately the same pressuredifferential as each other of said stages.